This invention relates to gas turbine engine shrouds and, more particularly, to a shroud having cooling passages that increase efficiency of the gas turbine engine.
Conventional gas turbine engines are widely known and used to propel aircraft and other vehicles. Typically, gas turbine engines include a compressor section, a combustor section, and a turbine section. Compressed air from the compressor section is fed to the combustor section and mixed with fuel. The combustor ignites the fuel and air mixture to produce a flow of hot gases. The turbine section transforms the flow of hot gases into mechanical energy to drive the compressor. An exhaust nozzle directs the hot gases out of the gas turbine engine to provide thrust to the aircraft or other vehicle.
Typically, shroud sections, also known as blade outer air seals, are located radially outward from the turbine section and function as an outer wall for the hot gas flow through the gas turbine engine. The shroud sections typically include a cooling system, such as a cast, cored, internal cooling passage, to maintain the shroud sections at a desirable temperature. Cooling air is forced through the cooling passages and bleeds into the hot gas flow.
Rotation of turbine blades relative to turbine vanes in the turbine section causes a circumferential component of hot gas flow relative to the engine axis. In conventional shroud sections, the cooling air bleeds into the hot gas flow along an axial direction. Disadvantageously, axial momentum of the discharged cooling air acts against circumferential momentum of the hot gas flow to undesirably reduce the overall momentum of the hot gas flow. This results in an aerodynamic disadvantage that reduces efficiency of turbine blade rotation.
Accordingly, there is a need for shroud sections having cooling passages that minimize momentum loss of the hot gas flow. This invention addresses these needs and provides enhanced capabilities while avoiding the shortcomings and drawbacks of the prior art.